The present invention relates generally to gas turbine engines, and, more specifically, to turbine efficiency therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the combustion gases in turbine stages which power the compressor through one drive shaft, and produce additional work for powering an upstream fan in a turbofan aircraft engine application, or driving an external drive shaft for marine and industrial (M&I) applications.
The basic core engine typically includes a multistage axial compressor having rows of compressor blades and corresponding guide vanes which pressurize ambient air in stages and correspondingly increase the temperature thereof. The air discharged from the aft end of the compressor has the highest pressure, commonly referred to as compressor discharge pressure (CDP), and a correspondingly high temperature.
In an exemplary configuration, the compressor may have seven stages for increasing air pressure many times atmospheric pressure along with many hundreds of degrees of temperature increase due to the compression cycle. A fewer or greater number of compression stages may be used as desired for the specific design of the gas turbine engine and its intended use.
A majority of the CDP air discharged from the compressor is mixed with fuel in the combustor for generating hot combustion gases. These combustion gases then undergo an expansion cycle in the several turbine stages for extracting energy therefrom which correspondingly reduces the pressure of the combustion gases and the temperature thereof. A high pressure turbine (HPT) immediately follows the combustor and is used to power the compressor blades in the core engine.
A low pressure turbine (LPT) follows the HPT and drives the second shaft for powering the upstream fan in the turbofan engine application, or driving an external drive shaft for M&I applications.
The overall efficiency of the gas turbine engine is dependent on the efficiency of air compression, efficiency of combustion, and efficiency of combustion gas expansion in the turbine stages.
Each turbine stage typically includes an upstream turbine nozzle or stator having a row of nozzle vanes which direct the combustion gases downstream through a corresponding row of turbine rotor blades. The blades are typically mounted to the perimeter of a supporting rotor disk in corresponding dovetail slots formed therein.
The turbine blades and vanes are typically hollow airfoils with corresponding internal cooling channels therein which receive compressor discharge air for cooling thereof during operation. The hollow blades and vanes typically include various rows of film cooling and other discharge holes through the pressure and suction sidewalls thereof for discharging the spent internal cooling air in corresponding external films for further protecting the airfoils.
The main turbine flowpath is designed to confine the combustion gases as they flow through the engine and decrease in temperature and pressure from the combustor. The various cooling circuits for the turbine components are independent from the main flowpath and must be provided with cooling air at sufficient pressure to prevent ingestion of the hot combustion gases therein during operation.
For example, suitable rotary seals are provided between the stationary turbine nozzles and the rotating turbine blades to prevent ingestion or backflow of the hot combustion gases into the cooling circuits.
Since the airfoils of the nozzle vanes and turbine blades typically include rows of cooling air outlet holes, the cooling air must also have sufficient pressure greater than that of the external combustion gases to provide a suitable backflow margin to prevent ingestion of the hot combustion gases into the turbine airfoils themselves.
Since the combustion gases and cooling air are channeled through corresponding flowpaths or flow circuits in the engine, they are subject to various aerodynamic losses which further decrease engine efficiency. Fluid flow is subject to friction or drag losses, flow separation losses, and mixing losses all of which reduce pressure and decrease efficiency.
The rotary seal between the first stage turbine nozzle and first stage turbine rotor blades is at one critical site which significantly affects turbine efficiency. The nozzle including its inner band is a stationary or stator component immediately followed downstream by the rotating turbine blades and their corresponding rotating platforms which form the radially inner flowpath boundary for the hot combustion gases being channeled through the first stage turbine.
The combustion gases discharged from the turbine nozzle must necessarily flow over the axial gap and rotor seal therebetween to reach the turbine blades. The rotary seal cooperates with the internal pressurized purge air being channeled through the axial gap to prevent backflow of the hot combustion gases into the purge cooling circuit.
Accordingly, the smooth flow of the hot combustion gases is interrupted at the inner band-platform axial gap, and the purge air mixes with the combustion gases at this site and further degrades smooth flow.
Flow separation of the combustion gases occurs at this rotary seal and the combustion gases mix with the different velocity purge air for collectively further reducing efficiency.
Efficiency losses at this location decrease the total pressure of the combustion gases available at the turbine blades, which in turn corresponding reduces turbine efficiency. And, the disrupted flow at the beginning of the blade platforms can lead to increased heating thereof.
Accordingly, it is desired to provide a turbine stage having an improved rotary seal between the stator nozzle and rotor blades for improving turbine efficiency.